1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an industrial gas turbine engine turbine blade with platform cooling.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
The cooling of the blade platform in an industrial gas turbine engine is produced using convection cooling or film cooling. In the convection cooled platform, straight cooling holes formed within the platform with long length-to-diameter ratios are used. FIGS. 1 and 2 show this prior art blade platform cooling design using film cooling holes. FIGS. 3 and 4 show the prior art blade platform cooling design using convection cooling holes. The blade includes an airfoil section extending from a platform 12 and a root section 13 with a cooling air supply channel 16. In FIG. 2, the platform is cooled using a number of film cooling holes 15 connected to a dead rim cavity 14 formed below the platform 12. In FIG. 4, the platform convection cooling holes 16 are supplied from the cooling air supply channel 17.
The blade platform cooling designs of FIGS. 1 through 4 have several important design issues. Providing film cooling air for the entire blade platform requires a cooling air supply pressure from the dead rim cavity 14 to be higher than the peak blade platform external gas side pressure. This design induces a high leakage flow around the blade attachment region 13 and therefore causes a performance penalty. Using the long length-to-diameter ratio convection cooling holes that are drilled from the platform edge to the airfoil cooling supply channel 16 from the blade platform produces unacceptable stress levels at the airfoil cooling core and the platform cooling channels interface location, which therefore yields a low blade life. This problem is primary due to the large mass at the front and back ends of the blade root or attachment 13 which constrains the blade platform expansion. The cooling channels are also oriented transverse to the primary direction of the stress field which produces high stress concentrations in the cooling channels at the entrance location.
An inspection of these prior art turbine rotor blades indicates that an over-temperature occurs at the blade platform pressure side location and the aft portion of the suction side platform edge and the aft section of the suction side platform to airfoil transition location. To address this over-temperature problem, the blade platform cooling circuit of FIG. 5 was proposed. FIG. 5 shows three convection cooling channels 23 with long length-to-diameter ratios that are parallel to the platform to cool the platform pressure side surface and a large diameter cooling channel 21 with three smaller channels 22 that branch off to cool the platform suction side surface. Cooling air form the four larger diameter cooling channels 21 and 23 are supplied from a front end of the platform from the dead rim cavity below the platform and then discharged at the aft face of the platform into a gap formed between adjacent platform edges.
In some turbine blade, the airfoil suction side surface is positioned very close to the mate-face of the platform and thus not enough space is available for a cooling air channel to extend from the leading edge to the trailing edge of the platform. Also, a long and straight cooling channel used on the pressure side of the platform cannot provide sufficient cooling for the platform hot spot locations.